|Publication number||US4671735 A|
|Application number||US 06/692,176|
|Publication date||Jun 9, 1987|
|Filing date||Jan 17, 1985|
|Priority date||Jan 19, 1984|
|Also published as||DE3401742A1, DE3401742C2|
|Publication number||06692176, 692176, US 4671735 A, US 4671735A, US-A-4671735, US4671735 A, US4671735A|
|Inventors||Axel Rossmann, Wilhelm Hoffmueller, Josef Eichner|
|Original Assignee||Mtu-Motoren-Und Turbinen-Union Munchen Gmbh|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (15), Referenced by (27), Classifications (9), Legal Events (6)|
|External Links: USPTO, USPTO Assignment, Espacenet|
In compressors, especially axial-flow compressors, efficiency and operating behavior depend to a large extent on the ability to keep the width of the radial clearance or gap between the rotor blades and the casing as small as possible. In order to minimize the radial clearance, it can be adjusted by abrasion caused by blades rubbing against a casing liner (abradable running-in coating). The blades of the turbomachine, however, should suffer as little abrasive wear as possible, because otherwise, and especially with ovalized casings, the gaps grow and remain unfavorably large and the blades must be repaired at great cost or must be discarded for being too short.
The abradable (running-in) coatings on the casing, especially when they are soft, cause little blade tip wear, but have been shown to be sensitive to erosion and temperature.
A soft abradable coating has been described, for example, in GB-PS No. 733,918.
Described in DE-OS No. 28 53 958 and its U.S. counterpart, U.S. Pat. No. 4,227,703, is a gas seal and method for its manufacture, in which the point of a protruding portion of a composite material is fixedly connected to a turbine blade and takes the shape of a knife edge or fin. These fins are also referred to as squealer tips. From these tips, abrasive particles project in the direction of the protrusion. In this construction of an abrasive tip, particles are knocked out of the opposite sealing member to an undesirable degree.
It is the object of the present invention to provide a simple and inexpensive means for adjusting, especially for minimizing the radial clearance between the rotor blades and the casing of a compressor, especially of an aircraft gas engine or gas turbine power plant, which while being tolerant of wear, nevertheless minimizes the same.
The underlying problems are solved according to the present invention with rotors of the type described above in that the rubbing work is done by only one or a few "rubbing blades" in each compressor stage of a rotor. In order to achieve this favorable rubbing action the blade, in its area facing the casing, takes the shape of a shroud, and the shroudlike end area of the blade carries on its radially outer side a conventional hardfacing (protective layer) which is matched to the abradable coating on the casing. Matching of the material of the hard facing on the blade with that of the abradable coating is essential to provide minimal and smooth wearing of the abradable coating. In matching the properties of the materials, one need keep in mind that minimal or no wear is desired from the hard surface on the blade and thus a hard material is desired, such that portions thereof will not be pealed off from the blade tip. Coeficients of expansion of the matched materials should be as close to one another as possible so that heat changes do not produce relative expansions that could cause excessive abraiding or excessive leakage around the blade tip. What is desired is a good contact of the abradable material on the blade tip and a smooth rubbing action by the blade tip, without excessive wear and without undesirable rubbing or tearing. Lastly, the matching should pair two materials that will not enter into undesirable reactions with one another and especially chemical reactions.
The major advantages of the present invention reside in that the new blade with its wear-resistant tip enables harder, erosion and heat resistant abradable coatings to be used on the casing without aggravating blade tip wear. The radial clearance between the rotor blades and the casing is easy to adjust and to minimize, owing to the shroud-like widening of the blade and to the selection of the layer thickness of hardfacing applied to the blade tip.
The inventive concept expressly embraces all combinations and subcombinations of the features contained herein, also when combined with known features.
These and other objects, features and advantages of the present invention will become more apparent from the following description when taken in connection with the accompanying drawing which shows, for purposes of illustration only, one embodiment in accordance with the present invention, and wherein:
FIG. 1 is a schematic elevational view of a rotor with individual blades in accordance with the present invention;
FIG. 2 is a view of a blade in accordance with the present invention, taken in a direction transverse to the direction of flow through an axial-flow compressor;
FIG. 3 is a perspective view of the blade of FIG. 2, taken essentially in the direction of flow, and
FIG. 4 is a plan view illustrating the blade of FIGS. 2 and 3.
Referring now to the drawing, wherein like reference numerals are used throughout the various views to designate corresponding parts, FIG. 1 illustrates a rotor with two symmetrically arranged blades in accordance with the present invention.
As it will become apparent from FIG. 2, the blades take a shape normal for rotor blades used in axial-flow compressors. The configuration, the material and the type of manufacture of the blade can be selected within wide limits. This also applies to the manner of securing the blade root in the rotor disk.
The wear-resistant layer 1 (hardfacing) is deposited on a widened portion in the end area 2 of the tip of the blade 3. The root is indicated by the numeral 4 and has an inner platform 5 (cf. especially FIG. 3). The contour of the twisted blade 3 is shown in FIG. 4 in broken line. It takes the form of an aerofoil section.
The hardfacing is deposited on the radially outer end of the blade, on the tip thereof and adjacent to the casing and especially adjacent to its abradable coating. The layer consists, at least on the surface facing the casing, of a wear-resistant material, such as a hard material. Suited especially for that purpose are materials of the tungsten carbide, silicon carbide, chromium carbide group in applications in the lower and medium temperature ranges. For elevated temperatures, the layer can advantageously be titanium carbide, titanium nitride or silicon nitride. Also other materials of comparable wear-resistance would be suitable for the intents of the present invention, especially ceramic materials, such as metal oxides or other metal compounds and also mixtures of materials. When selecting the material for the hardfacing, however, it should be remembered that it should bond readily with the blade material, such as steel, nickel, chromium, titanium alloys or others, and that it should mate well with the abradable coating on the casing, against which it may rub. Excessive wear and undesirable rubbing should be avoided especially when the casing becomes out-of-round to take an oval or polygonal shape. This may happen in transient operating states, such as start, acceleration and shut-down or coast-down of the axial-flow compressor. These conditions may give rise to irregular thermal and/or mechanical loads and stresses and non-uniform expansions of the casing and rotor.
If the wear-resistant layer of the present invention is selected from among such materials as will optimally reduce abrasive wear, no risk is involved at the blade tip. With a suitable selection of matched or paired materials to take into consideration, the problems mentioned above the radial clearance between the rotor blade tips and the casing can also be kept practically constant. Nor should the mated materials of the hardfacing on a blade tip and the abradable coating on the casing enter into undesirable reactions, and especially chemical reactions should be avoided.
The above-mentioned preferred materials are deposited on the blade tip at its shroud-like widened portion in the end area 2 either directly or with the intervention of a bond layer, for example, by detonation or plasma spray process or by physical or chemical vapor deposition (PVD or CVD). The process should be chosen to suit the materials selected for the layer taking into consideration the above-mentioned conditions. The preferred layer thicknesses thereby lie between about 0.1 mm to about 1 mm, however, they may conventionally deviate either way from this range depending on the process and material selected.
A blade provided with the wear-resistant layer 1 of the present invention is bound to be appreciably heavier than the other blades of the same stage that do not have the layer and the shroud-like widening at the blade tip. The extra weight will lead to greater centrifugal forces and, thus, to increased low-cycle fatigue in the blade root. If necessary, this situation can be remedied, however, by giving the blade root larger dimensions than the root of other blades. The corresponding slot in the rotor disk will likewise have to be adjusted to accommodate the changed size of the blade root. Rather than changing dimensions, however, it may be preferable to select a different material for the root 4, which should then be fixedly connected to the airfoil 3, e.c., by welding.
If it is intended to install more than one blade of the present invention in the rotor disk, use should preferably be made of an even number of blades, and the blades should be symmetrically spaced around the circumference of the disk to prevent imbalance or other problems (FIG. 1).
The inventive concept naturally embraces also modifications and other versions. To be conducive to favorable flow conditions, the width of the shroud need not necessarily bridge the full distance between blades.
The invention finds preferred use in axial-flow compressors of aircraft engines in combination with gas turbines, where normally several compressor stages (also alternating with guide vanes) plus several turbine stages are arranged on a shaft. In this arrangement, the thermal load on the compressor blades is as a rule less than that on the hot gas wetted turbine blades.
While we have shown and described only one embodiment in accordance with the present invention, it is understood that the same is not limited thereto but is susceptible of numerous changes and modifications as known to those skilled in the art, and we therefore do not wish to be limited to the details shown and described herein, but intend to cover all such changes and modifications as are encompassed by the scope of the appended claims.
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US941395 *||May 2, 1905||Nov 30, 1909||Westinghouse Machine Co||Elastic-fluid turbine.|
|US1360936 *||May 13, 1919||Nov 30, 1920||British Westinghouse Electric||Fluid-pressure turbine|
|US2023111 *||Sep 8, 1934||Dec 3, 1935||Westinghouse Electric & Mfg Co||Silent fan|
|US3199836 *||May 4, 1964||Aug 10, 1965||Gen Electric||Axial flow turbo-machine blade with abrasive tip|
|US3537713 *||Feb 21, 1968||Nov 3, 1970||Garrett Corp||Wear-resistant labyrinth seal|
|US3975165 *||Dec 26, 1973||Aug 17, 1976||Union Carbide Corporation||Graded metal-to-ceramic structure for high temperature abradable seal applications and a method of producing said|
|US4148494 *||Dec 21, 1977||Apr 10, 1979||General Electric Company||Rotary labyrinth seal member|
|US4218066 *||Mar 23, 1976||Aug 19, 1980||United Technologies Corporation||Rotary seal|
|US4227703 *||Nov 27, 1978||Oct 14, 1980||General Electric Company||Gas seal with tip of abrasive particles|
|US4390320 *||May 1, 1980||Jun 28, 1983||General Electric Company||Tip cap for a rotor blade and method of replacement|
|US4411589 *||Oct 21, 1980||Oct 25, 1983||Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A."||Retaining device for the compressor casing of a turbine engine|
|US4466785 *||Nov 18, 1982||Aug 21, 1984||Ingersoll-Rand Company||Clearance-controlling means comprising abradable layer and abrasive layer|
|US4477226 *||May 9, 1983||Oct 16, 1984||General Electric Company||Balance for rotating member|
|GB733918A *||Title not available|
|JPS4739867A *||Title not available|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US4808055 *||Apr 15, 1987||Feb 28, 1989||Metallurgical Industries, Inc.||Turbine blade with restored tip|
|US4874290 *||Aug 26, 1988||Oct 17, 1989||Solar Turbines Incorporated||Turbine blade top clearance control system|
|US5292382 *||Sep 5, 1991||Mar 8, 1994||Sulzer Plasma Technik||Molybdenum-iron thermal sprayable alloy powders|
|US5530050 *||Apr 6, 1994||Jun 25, 1996||Sulzer Plasma Technik, Inc.||Thermal spray abradable powder for very high temperature applications|
|US6478304 *||Jul 14, 2000||Nov 12, 2002||Mtu Aero Engines Gmbh||Sealing ring for non-hermetic fluid seals|
|US6688867||Oct 4, 2001||Feb 10, 2004||Eaton Corporation||Rotary blower with an abradable coating|
|US6984107 *||Jan 27, 2003||Jan 10, 2006||Mtu Aero Engines Gmbh||Turbine blade for the impeller of a gas-turbine engine|
|US7425115||Oct 14, 2005||Sep 16, 2008||Alstom Technology Ltd||Thermal turbomachine|
|US8708655||Sep 24, 2010||Apr 29, 2014||United Technologies Corporation||Blade for a gas turbine engine|
|US9021696||Apr 21, 2010||May 5, 2015||MTU Aero Engines AG||Method for producing a plating of a vane tip and correspondingly produced vanes and gas turbines|
|US20030170120 *||Jan 27, 2003||Sep 11, 2003||Richard Grunke||Turbine blade for the impeller of a gas-turbine engine|
|US20110086163 *||Sep 30, 2010||Apr 14, 2011||Walbar Inc.||Method for producing a crack-free abradable coating with enhanced adhesion|
|US20130078084 *||Sep 23, 2011||Mar 28, 2013||United Technologies Corporation||Airfoil air seal assembly|
|US20150093237 *||Sep 30, 2013||Apr 2, 2015||General Electric Company||Ceramic matrix composite component, turbine system and fabrication process|
|EP1331362A2 *||Jan 27, 2003||Jul 30, 2003||Kabushiki Kaisha Toshiba||Geothermal steam turbine|
|EP1462617A2 *||Mar 16, 2004||Sep 29, 2004||ALSTOM Technology Ltd||Blading for an axial-flow turbomachine|
|EP1803898A2 *||Jan 27, 2003||Jul 4, 2007||Kabushiki Kaisha Toshiba||Geothermal turbine|
|EP1803898A3 *||Jan 27, 2003||Jan 4, 2012||Kabushiki Kaisha Toshiba||Geothermal turbine|
|EP2309097A1||Sep 30, 2009||Apr 13, 2011||Siemens Aktiengesellschaft||Airfoil and corresponding guide vane, blade, gas turbine and turbomachine|
|EP2309098A1||Sep 30, 2009||Apr 13, 2011||Siemens Aktiengesellschaft||Airfoil and corresponding guide vane, blade, gas turbine and turbomachine|
|EP2573326A1 *||Aug 31, 2012||Mar 27, 2013||United Technologies Corporation||Airfoil tip air seal assembly|
|WO2001006097A1 *||Jul 14, 2000||Jan 25, 2001||MTU MOTOREN- UND TURBINEN-UNION MüNCHEN GMBH||Sealing ring for non-hermetic fluid seals|
|WO2003064818A2 *||Jan 28, 2003||Aug 7, 2003||Kabushiki Kaisha Toshiba||Geothermal steam turbine|
|WO2003064818A3 *||Jan 28, 2003||Nov 27, 2003||Toshiba Kk||Geothermal steam turbine|
|WO2010121597A3 *||Apr 21, 2010||Jul 7, 2011||Mtu Aero Engines Gmbh||Method for producing a plating of a vane tip of a gas turbine blade|
|WO2011038966A1||Aug 9, 2010||Apr 7, 2011||Siemens Aktiengesellschaft||Airfoil and corresponding guide vane, blade, gas turbine and turbomaschine|
|WO2011038971A1||Aug 10, 2010||Apr 7, 2011||Siemens Aktiengesellschaft||Airfoil and corresponding guide vane, blade, gas turbine and turbomaschine|
|U.S. Classification||415/173.1, 415/173.4|
|International Classification||F04D29/16, F04D29/38, F01D5/20|
|Cooperative Classification||Y02T50/672, Y02T50/673, F01D5/20|
|Jan 17, 1985||AS||Assignment|
Owner name: MTU-MOTOREN-UND TURBINEN-UNION MUNCHEN GMBH MUNICH
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST.;ASSIGNORS:ROSSMANN, AXEL;HOFFMUELLER, WILHELM;EICHNER, JOSEF;REEL/FRAME:004361/0930
Effective date: 19850103
|Aug 20, 1990||FPAY||Fee payment|
Year of fee payment: 4
|Nov 23, 1994||FPAY||Fee payment|
Year of fee payment: 8
|Dec 29, 1998||REMI||Maintenance fee reminder mailed|
|Jun 6, 1999||LAPS||Lapse for failure to pay maintenance fees|
|Aug 3, 1999||FP||Expired due to failure to pay maintenance fee|
Effective date: 19990609